U.S. patent application number 14/672380 was filed with the patent office on 2016-10-06 for hybrid nozzle segment assemblies for a gas turbine engine.
The applicant listed for this patent is General Electric Company. Invention is credited to Matthew Mark WEAVER.
Application Number | 20160290147 14/672380 |
Document ID | / |
Family ID | 55642296 |
Filed Date | 2016-10-06 |
United States Patent
Application |
20160290147 |
Kind Code |
A1 |
WEAVER; Matthew Mark |
October 6, 2016 |
HYBRID NOZZLE SEGMENT ASSEMBLIES FOR A GAS TURBINE ENGINE
Abstract
A nozzle segment assembly for a gas turbine engine may generally
include inner and outer ring support segments and a nozzle fairing
positioned between the inner and outer ring support segments. The
nozzle fairing may be formed from a ceramic matrix composite (CMC)
material and may include both an outer endwall and an inner
endwall. In addition, the nozzle fairing may include a strut vane
extending between the inner and outer endwalls. The nozzle segment
assembly may also include a metallic strut extending through the
strut vane between the outer and inner ring supports and at least
one secondary vane configured to be received through at least one
of the outer endwall or the inner endwall of the nozzle fairing
such that the at least one secondary vane extends between the inner
and outer endwalls at a location adjacent to the strut vane.
Inventors: |
WEAVER; Matthew Mark;
(Loveland, OH) |
|
Applicant: |
Name |
City |
State |
Country |
Type |
General Electric Company |
Schenectady |
NY |
US |
|
|
Family ID: |
55642296 |
Appl. No.: |
14/672380 |
Filed: |
March 30, 2015 |
Current U.S.
Class: |
1/1 |
Current CPC
Class: |
F05D 2300/175 20130101;
F05D 2220/32 20130101; Y02T 50/60 20130101; F01D 25/005 20130101;
F01D 9/047 20130101; F02C 3/04 20130101; F05D 2240/128 20130101;
Y02T 50/673 20130101; F01D 9/042 20130101; Y02T 50/672 20130101;
F05D 2300/6033 20130101; F01D 9/041 20130101; F02C 7/20 20130101;
F05D 2260/941 20130101; F05D 2300/607 20130101; F05D 2240/35
20130101 |
International
Class: |
F01D 9/04 20060101
F01D009/04; F02C 3/04 20060101 F02C003/04; F01D 25/00 20060101
F01D025/00 |
Claims
1. A nozzle segment assembly for a gas turbine engine, the nozzle
segment assembly comprising: an outer ring support segment and an
inner ring support segment; a nozzle fairing positioned between the
inner and outer ring support segments, the nozzle fairing being
formed from a ceramic matrix composite (CMC) material, the nozzle
fairing including an outer endwall configured to be positioned
adjacent to the outer ring support segment and an inner endwall
configured to be positioned adjacent to the inner ring support
segment, the nozzle fairing further comprising a strut vane
extending between the inner and outer endwalls; a metallic strut
extending through the strut vane between the outer and inner ring
supports; and at least one secondary vane configured to be received
through at least one of the outer endwall or the inner endwall of
the nozzle fairing such that the at least one secondary vane
extends between the inner and outer endwalls at a location adjacent
to the strut vane.
2. The nozzle segment assembly of claim 1, wherein the outer
endwall defines an outer slot and the inner endwall defines an
inner slot, wherein separate portions of the at least one secondary
vane are configured to be received within the outer and inner
slots.
3. The nozzle segment assembly of claim 2, wherein the at least one
secondary vane extends radially between an outer vane end and an
inner vane end, the at least one secondary vane being configured to
extend through the inner and outer slots such that the outer vane
end extends radially outwardly from the outer endwall and the inner
vane end extend radially inwardly from the inner endwall.
4. The nozzle segment assembly of claim 3, wherein the outer vane
end is configured to be received within a vane recess defined in
the outer ring support.
5. The nozzle segment assembly of claim 4, wherein the at least one
secondary vane comprises a mounting tab extending from the outer
vane end, the mounting tab defining a through-hole configured to
receive a mounting pin for coupling the at least one secondary vane
to the outer ring support.
6. The nozzle segment assembly of claim 3, wherein the inner vane
end is configured to be received within a vane recess defined in
the inner ring support.
7. The nozzle segment assembly of claim 6, wherein the inner vane
end is configured to be coupled within the recess.
8. The nozzle segment assembly of claim 1, wherein the inner and
outer endwalls and the strut vane are formed integrally as a single
unitary component.
9. The nozzle segment assembly of claim 1, wherein the metallic
strut extends radially inwardly from the outer ring support through
the strut vane.
10. The nozzle segment assembly of claim 9, wherein the outer ring
support segment comprises an outer band segment, the metallic strut
and the outer band segment being formed integrally as a single
unitary component.
11. The nozzle segment assembly of claim 9, wherein the metallic
strut extends radially inwardly from the outer ring support to a
tip end, the tip end being configured to be received within a strut
recess defined in the inner ring support segment.
12. The nozzle segment assembly of claim 1, wherein a gap is
defined between an inner surface of the strut vane and an outer
surface of the strut when the strut is received within the strut
vane.
13. The nozzle segment assembly of claim 12, wherein the gap is
configured to receive a cooling medium for cooling at least one of
the strut vane or the strut.
14. A gas turbine engine, comprising: a compressor; a combustor in
flow communication with the compressor; and a turbine configured to
receive combustion products from the combustor, the turbine
including a turbine nozzle having an annular array of nozzle
segment assemblies, each of the nozzle segment assemblies
comprising: an outer ring support segment and an inner ring support
segment; a nozzle fairing positioned between the inner and outer
ring support segments, the nozzle fairing being formed from a
ceramic matrix composite (CMC) material, the nozzle fairing
including an outer endwall configured to be positioned adjacent to
the outer ring support segment and an inner endwall configured to
be positioned adjacent to the inner ring support segment, the
nozzle fairing further comprising a strut vane extending between
the inner and outer endwalls; a metallic strut extending through
the strut vane between the outer and inner ring supports; and at
least one secondary vane configured to be received through at least
one of the outer endwall or the inner endwall of the nozzle fairing
such that the at least one secondary vane extends between the inner
and outer endwalls at a location adjacent to the strut vane.
15. The gas turbine engine of claim 14, wherein the outer endwall
defines an outer slot and the inner endwall defines an inner slot,
wherein separate portions of the at least one secondary vane are
configured to be received within the outer and inner slots.
16. The gas turbine engine of claim 15, wherein the at least one
secondary vane extends radially between an outer vane end and an
inner vane end, the at least one secondary vane being configured to
extend through the inner and outer slots such that the outer vane
end extends radially outwardly from the outer endwall and the inner
vane end extend radially inwardly from the inner endwall.
17. The gas turbine engine of claim 16, wherein the outer vane end
is configured to be received within a vane recess defined in the
outer ring support, the at least one secondary vane comprising a
mounting tab extending from the outer vane end, the mounting tab
defining a through-hole configured to receive a mounting pin for
coupling the at least one secondary vane to the outer ring
support.
18. The gas turbine engine of claim 16, wherein the inner vane end
is configured to be received within a vane recess defined in the
inner ring support.
19. The gas turbine engine of claim 14, wherein the inner and outer
endwalls and the strut vane are formed integrally as a single
unitary component.
20. The gas turbine engine of claim 14, wherein the metallic strut
extends radially inwardly from the outer ring support through the
strut vane to a tip end, the tip end being configured to be
received within a strut recess defined in the inner ring support
segment.
Description
FIELD OF THE INVENTION
[0001] The present subject matter relates generally to gas turbine
engines and, more particularly, to a hybrid nozzle segment assembly
for a gas turbine engine.
BACKGROUND OF THE INVENTION
[0002] A gas turbine engine includes a turbomachinery core having a
high pressure compressor, combustor, and high pressure turbine
("HPT") in serial flow relationship. The core is operable in a
known manner to generate a primary gas flow. The high pressure
turbine includes annular arrays ("rows") of stationary vanes or
nozzles that direct the gases exiting the combustor into rotating
blades or buckets. Collectively one row of nozzles and one row of
blades make up a "stage." Typically two or more stages are used in
serial flow relationship. These components operate in an extremely
high temperature environment and, thus, must often be cooled by air
flow to ensure adequate service life.
[0003] Due to the operating temperatures within the gas turbine
engine, it is desirable to utilize materials with low coefficients
of thermal expansion. For example, to operate effectively in such
strenuous temperature and pressure conditions, composite materials
have been developed, such as ceramic matrix composite (CMC)
materials. CMC materials provide both temperature and density
advantages over metallic materials, thereby making the materials
desirable options for manufacturing high temperature, hot gas path
components. However, CMC materials also have unique mechanical
properties that must be considered during the design and
application of such materials within the interior of a gas turbine
engine. For example, CMC materials have relatively low tensile
ductility or low strain to failure as compared to metallic
materials. As a result, CMC-based components are often not equipped
to handle significant mechanical loading during operation of the
gas turbine engine.
[0004] In this regard, attempts have been made to form nozzle vanes
from CMC materials to increase the high temperature capabilities of
such components. However, these prior attempts have not fully
addressed the mechanical loading issues for the nozzle vanes,
thereby leading to durability challenges for the CMC-based
components.
[0005] Accordingly, an improved nozzle segment assembly that
provides increased mechanical support for CMC-based nozzle vanes
would be welcomed in the technology.
BRIEF DESCRIPTION OF THE INVENTION
[0006] Aspects and advantages of the invention will be set forth in
part in the following description, or may be obvious from the
description, or may be learned through practice of the
invention.
[0007] In one aspect, the present subject matter is directed to a
nozzle segment assembly for a gas turbine engine. The nozzle
segment assembly may generally include an outer ring support
segment and an inner ring support segment. The nozzle segment
assembly may include a nozzle fairing positioned between the inner
and outer ring support segments. The nozzle fairing may be formed
from a ceramic matrix composite (CMC) material and may include both
an outer endwall configured to be positioned adjacent to the outer
ring support segment and an inner endwall configured to be
positioned adjacent to the inner ring support segment. In addition,
the nozzle fairing may include a strut vane extending between the
inner and outer endwalls. The nozzle segment assembly may also
include a metallic strut extending through the strut vane between
the outer and inner ring supports and at least one secondary vane
configured to be received through at least one of the outer endwall
or the inner endwall of the nozzle fairing such that the at least
one secondary vane extends between the inner and outer endwalls at
a location adjacent to the strut vane.
[0008] In another aspect, the present subject matter is directed to
a gas turbine engine. The gas turbine engine may generally include
a compressor, a combustor in flow communication with the compressor
and a turbine configured to receive combustion products from the
combustor. The turbine may include a turbine nozzle having an
annular array of nozzle segment assemblies. Each nozzle segment
assembly may generally include an outer ring support segment, an
inner ring support segment and a nozzle fairing positioned between
the inner and outer ring support segments. The nozzle fairing may
be formed from a ceramic matrix composite (CMC) material and may
include both an outer endwall configured to be positioned adjacent
to the outer ring support segment and an inner endwall configured
to be positioned adjacent to the inner ring support segment. In
addition, the nozzle fairing may include a strut vane extending
between the inner and outer endwalls. Each nozzle segment assembly
may also include a metallic strut extending through the strut vane
between the outer and inner ring supports and at least one
secondary vane configured to be received through at least one of
the outer endwall or the inner endwall of the nozzle fairing such
that the at least one secondary vane extends between the inner and
outer endwalls at a location adjacent to the strut vane.
[0009] These and other features, aspects and advantages of the
present invention will become better understood with reference to
the following description and appended claims. The accompanying
drawings, which are incorporated in and constitute a part of this
specification, illustrate embodiments of the invention and,
together with the description, serve to explain the principles of
the invention.
BRIEF DESCRIPTION OF THE DRAWINGS
[0010] A full and enabling disclosure of the present invention,
including the best mode thereof, directed to one of ordinary skill
in the art, is set forth in the specification, which makes
reference to the appended figures, in which:
[0011] FIG. 1 illustrates a cross-sectional view of one embodiment
of a gas turbine engine that may be utilized within an aircraft in
accordance with aspects of the present subject matter;
[0012] FIG. 2 illustrates a perspective view of one embodiment of a
turbine nozzle suitable for use within the gas turbine shown in
FIG. 1 in accordance with aspects of the present subject
matter;
[0013] FIG. 3 illustrates a perspective view of one embodiment of a
nozzle segment assembly suitable for use within the turbine nozzle
shown in FIG. 2 in accordance with aspects of the present subject
matter;
[0014] FIG. 4 illustrates an exploded view of the nozzle segment
assembly shown in FIGS. 3; and
[0015] FIG. 5 illustrates a cross-sectional view of the nozzle
segment assembly shown in FIG. 3 taken about line 5-5.
DETAILED DESCRIPTION OF THE INVENTION
[0016] Reference now will be made in detail to embodiments of the
invention, one or more examples of which are illustrated in the
drawings. Each example is provided by way of explanation of the
invention, not limitation of the invention. In fact, it will be
apparent to those skilled in the art that various modifications and
variations can be made in the present invention without departing
from the scope or spirit of the invention. For instance, features
illustrated or described as part of one embodiment can be used with
another embodiment to yield a still further embodiment. Thus, it is
intended that the present invention covers such modifications and
variations as come within the scope of the appended claims and
their equivalents.
[0017] In general, the present subject matter is directed to a
nozzle segment assembly forming part of a turbine nozzle for a gas
turbine engine. Specifically, in several embodiments, the nozzle
segment assembly may include a hybrid CMC/metal vane design,
wherein a structural metallic component or strut is shielded from
the combustion gases flowing along the hot gas path of the gas
turbine engine by a CMC fairing and corresponding CMC endwalls. In
addition, the nozzle segment assembly may also include one or more
secondary CMC-based vanes configured to be simply supported
along-side the CMC shielded strut. For example, in one embodiment,
the nozzle segment assembly may correspond to a triplet vane
design. In such an embodiment, the nozzle segment assembly may
include two secondary CMC-based vanes corresponding to the outer
vanes of the assembly, with the CMC fairing serving as the center
vane.
[0018] It should be appreciated that the disclosed hybrid CMC/metal
design may allow for a nozzle segment assembly to be provided the
advantages of both CMC- and metallic-based components. For example,
the metallic strut may serve to increase the mechanical loading
capability of the assembly. In addition, the high temperature
capabilities of the various CMC components may serve to reduce the
cooling flow requirements for the assembly. Moreover, given their
significantly lower densities, the CMC-based components may also
serve to reduce the overall weight of the nozzle segment assembly
when compared to conventional metallic-based nozzle components.
[0019] Referring now to the drawings, FIG. 1 illustrates a
cross-sectional view of one embodiment of a gas turbine engine 10
that may be utilized within an aircraft in accordance with aspects
of the present subject matter, with the engine 10 being shown
having a longitudinal or axial centerline axis 12 extending
therethrough for reference purposes. In general, the engine 10 may
include a core gas turbine engine (indicated generally by reference
character 14) and a fan section 16 positioned upstream thereof. The
core engine 14 may generally include a substantially tubular outer
casing 18 that defines an annular inlet 20. In addition, the outer
casing 18 may further enclose and support a booster compressor 22
for increasing the pressure of the air that enters the core engine
14 to a first pressure level. A high pressure, multi-stage,
axial-flow compressor 24 may then receive the pressurized air from
the booster compressor 22 and further increase the pressure of such
air. The pressurized air exiting the high-pressure compressor 24
may then flow to a combustor 26 within which fuel is injected into
the flow of pressurized air, with the resulting mixture being
combusted within the combustor 26. The high energy combustion
products are directed from the combustor 26 along the hot gas path
of the engine 10 to a first (high pressure) turbine 28 for driving
the high pressure compressor 24 via a first (high pressure) drive
shaft 30, and then to a second (low pressure) turbine 32 for
driving the booster compressor 22 and fan section 16 via a second
(low pressure) drive shaft 34 that is generally coaxial with first
drive shaft 30. After driving each of turbines 28 and 32, the
combustion products may be expelled from the core engine 14 via an
exhaust nozzle 36 to provide propulsive jet thrust.
[0020] It should be appreciated that each turbine 28, 30 may
generally include one or more turbine stages, with each stage
including a turbine nozzle (not shown in FIG. 1) and a downstream
turbine rotor (not shown in FIG. 1). As will be described below,
the turbine nozzle may include a plurality of vanes disposed in an
annular array about the centerline axis 12 of the engine 10 for
turning or otherwise directing the flow of combustion products
through the turbine stage towards a corresponding annular array of
rotor blades forming part of the turbine rotor. As is generally
understood, the rotor blades may be coupled to a rotor disk of the
turbine rotor, which is, in turn, rotationally coupled to the
turbine's drive shaft (e.g., drive shaft 30 or 34).
[0021] Additionally, as shown in FIG. 1, the fan section 16 of the
engine 10 may generally include a rotatable, axial-flow fan rotor
38 that configured to be surrounded by an annular fan casing 40. It
should be appreciated by those of ordinary skill in the art that
the fan casing 40 may be configured to be supported relative to the
core engine 14 by a plurality of substantially radially-extending,
circumferentially-spaced outlet guide vanes 42. As such, the fan
casing 40 may enclose the fan rotor 38 and its corresponding fan
rotor blades 44. Moreover, a downstream section 46 of the fan
casing 40 may extend over an outer portion of the core engine 14 so
as to define a secondary, or by-pass, airflow conduit 48 that
provides additional propulsive jet thrust.
[0022] During operation of the engine 10, it should be appreciated
that an initial air flow (indicated by arrow 50) may enter the
engine 10 through an associated inlet 52 of the fan casing 40. The
air flow 50 then passes through the fan blades 44 and splits into a
first compressed air flow (indicated by arrow 54) that moves
through conduit 48 and a second compressed air flow (indicated by
arrow 56) which enters the booster compressor 22. The pressure of
the second compressed air flow 56 is then increased and enters the
high pressure compressor 24 (as indicated by arrow 58). After
mixing with fuel and being combusted within the combustor 26, the
combustion products 60 exit the combustor 26 and flow through the
first turbine 28. Thereafter, the combustion products 60 flow
through the second turbine 32 and exit the exhaust nozzle 36 to
provide thrust for the engine 10.
[0023] Referring now to FIG. 2, a perspective view of one
embodiment of a turbine nozzle 100 is illustrated in accordance
with aspects of the present subject matter. The turbine nozzle 100
may generally be configured to be positioned within any suitable
turbine section of a gas turbine engine, such as the first (or high
pressure) turbine 28 and/or the second (or lower pressure) turbine
32 of the gas turbine engine 10 described above with reference to
FIG. 1.
[0024] In general, the turbine nozzle 100 may define a ring-like
shape formed by an annular array of nozzle segment assemblies 102.
As is generally understood, the nozzle segment assemblies 102 may
be configured to direct the combustion gases flowing along the hot
gas path of the turbine engine 10 downstream through a subsequent
row of rotor blades (not shown) extending radially outwardly from a
supporting rotor disk. As shown in FIG. 2, each nozzle segment
assembly 102 may include a nozzle fairing 104 having a strut vane
106 extending between radially inner and outer endwalls 108, 110
and one or more secondary nozzle vanes 112 supported between the
endwalls 108, 110. Additionally, each nozzle segment assembly 102
may include an inner support ring segment 114 and a hanger or outer
support ring segment 116. As shown in FIG. 2, the inner and outer
support ring segments 114, 116 may be arranged in an annular array
(e.g., about the engine centerline axis 12) so as to define the
inner and outer ring-shaped perimeters, respectively, of the
turbine nozzle 100. As will be described below, the outer support
ring segment 116 of each nozzle segment assembly 102 may generally
include a strut (not shown in FIG. 2) extending radially inwardly
therefrom that is configured to be received within the strut vane
106 so as to provide structural and load-bearing support for the
nozzle segment assembly 102.
[0025] Referring now to FIGS, 3-5, several views of one of the
nozzle segment assemblies 102 described above with reference to
FIG. 2 are illustrated in accordance with aspects of the present
subject matter. Specifically, FIG. 3 illustrates a perspective view
of the nozzle segment assembly 102 and FIG. 4 illustrates an
exploded of the nozzle segment assembly 102 shown in FIG. 4.
Additionally, FIG. 5 illustrates a cross-sectional view of the
nozzle segment assembly 102 shown in FIG. 3 taken about line
5-5.
[0026] As shown, the nozzle segment assembly 102 may include a
nozzle fairing 104 configured to be positioned between an inner
ring support segment 114 and an outer ring support segment 116. In
general, the nozzle fairing 104 may include a strut vane 106
extending radially between an inner endwall 108 and an outer
endwall 110. The strut vane 106 may be configured to define an
aerodynamic cross-sectional profile, such as by defining any
suitable airfoil-shape typically utilized for nozzle vanes within a
gas turbine engine. For example, as shown in FIG. 5, the strut vane
106 may define an airfoil shape including a pressure side 120 and a
suction side 122 extending between and a leading edge 124 and a
trailing edge 126. Additionally, as shown in FIG. 5, the strut vane
106 may be configured to be hollow. As will be described below,
such as a hollow configuration may allow for the vane 106 to
receive a load-bearing strut of the nozzle segment assembly
102.
[0027] Moreover, as particularly shown in FIG. 4, the strut vane
106 may be configured to extend radially end-to-end between the
inner and outer endwalls 108, 110. As a result, the inner and outer
endwalls 108, 110 may generally be configured to define radially
inner and outer flow path boundaries, respectively, for the hot
gases of combustion flowing along the hot gas path of the turbine
engine 10. For example, the combustion gases flowing past the strut
vane 106 may generally be directed between inner and outer endwalls
108, 110 along the aerodynamic shape of the vane 106.
[0028] In several embodiments, the strut vane 106 and the endwalls
108, 110 may be configured to be formed as a single, unitary
component. In such embodiments, the strut vane 106 and the endwalls
108, 110 may be formed integrally using any suitable manufacturing
process known in the art. As indicated above, such components may,
in one embodiment, be formed from a suitable CMC material. In such
instance, any suitable process known for manufacturing components
using CMC materials may be utilized to form the strut vane 106 and
the endwalls 108, 110 as an integral component, such as injection
molding, slip casting, tape casting, infiltration methods (e.g.,
chemical vapor infiltration, melt infiltration and/or the like) and
various other suitable methods and/or processes. However, in an
alternative embodiment, the strut vane 106 and the endwalls 108,
110 may be manufactured as separate components.
[0029] Additionally, as indicated above, the nozzle segment
assembly 102 may also include one or more secondary vanes 112
configured to be supported between the inner and outer endwalls
108, 110. For example, in the illustrated embodiment, the nozzle
segment assembly 102 is configured as a "triplet" design and, thus,
includes first and second secondary vanes 112 configured to extend
between the endwalls 108, 110 along opposite sides of the strut
vane 106. However, in another embodiment, the nozzle segment
assembly 102 may be configured as a "doublet" design. In such an
embodiment, the nozzle segment assembly 102 may only include a
single secondary vane 112 extending between the endwalls 108, 110
alongside the strut vane 106 (e.g., along one of the sides of the
strut vane 106).
[0030] As shown in the illustrated embodiment, each secondary vane
112 may generally extend radially between an inner vane end 130
(FIG. 4) and an outer vane end 132 (FIG. 4) and may define an
aerodynamic cross-sectional profile between its inner and outer
ends 130, 132. For example, as shown in FIG. 5, the secondary vanes
112 may define the same or a similar aerodynamic shape as the strut
vane 106, such as by defining a an airfoil-shape including a
pressure side 134 and a suction side 136 extending between a
leading edge 138 and a trailing edge 140. Additionally, in several
embodiments, each secondary vane 112 may be configured to be hollow
or substantially hollow so as to define one or more flow paths for
providing a cooling medium (e.g., air) through the vane 112. For
example, as shown in FIG. 5, each secondary vane 112 simply defines
a single interior cavity 142 for receiving a cooling medium.
However, in other embodiments, the secondary vanes 112 may define a
plurality of interior channels for receiving a cooling medium
(e.g., in a serpentine-like flow pattern within the interior of
each vane 112). Alternatively, the secondary vanes 112 may be
configured to define a solid cross-section.
[0031] When installing the secondary vanes 112 within the nozzle
segment assembly 102, the vanes 112 may, in several embodiments, be
configured to be inserted through one or both of the endwalls 108,
110. In this regard, the endwalls 108, 110 may define
airfoil-shaped slots (or other suitably shaped slots) for receiving
portions of the secondary vanes 112. For example, as shown in FIG.
4, the outer endwall 110 may define first and second outer slots
144, 146 for receiving corresponding portions of the first and
second vanes 112, respectively, and the inner sidewall 108 may
define first and second inner slots 148, 150 for receiving
corresponding portions of the first and second vanes 112,
respectively. In such an embodiment, the secondary vanes 112 may be
positioned within the slots 144, 146, 148, 150 such that the inner
vane end 130 of each vane 112 extends radially inwardly beyond the
inner endwall 108 and the outer vane end 132 of each vane 112
extends radially outwardly beyond the outer endwall 110. As will be
described below, by configuring the ends 130, 132 of the secondary
vanes 112 to extend radially beyond the inner and outer endwalls
108, 110, the secondary vanes 112 may be configured to be attached
to or secured within portions of the inner and outer ring support
segments 114 116 to ensure that the vanes 112 are properly
supported within the nozzle segment assembly 102.
[0032] It should be appreciated that, in several embodiments, the
strut vane 106, the endwalls 108, 110 and the secondary vanes 112
may all be formed from a non-metallic material having a relatively
low coefficient of thermal expansion. For instance, as indicated
above, the strut vane 106, the endwalls 108, 110 and the secondary
vanes 112 may all be formed from a CMC material. In such
embodiments, the CMC material may generally correspond to any
suitable CMC material known in the art and, thus, may generally
include a ceramic matrix having a suitable reinforcing material
incorporated therein to enhance the material's properties (e.g.,
the material strength and/or the thermo-physical properties). In
one embodiment, the CMC material used may be configured as a
continuous fiber reinforced CMC material. For example, suitable
continuous fiber reinforced CMC materials may include, but are not
limited to, CMC materials reinforced with continuous carbon fibers,
oxide fibers, silicon carbide monofilament fibers and other CMC
materials including continuous fiber lay-ups and/or woven fiber
preforms. In other embodiments, the CMC material used may be
configured as a discontinuous reinforced CMC material. For
instance, suitable discontinuous reinforced CMC materials may
include, but are not limited to, particulate, platelet, whisker,
discontinuous fiber, in situ and nano-composite reinforced CMC
materials.
[0033] As indicated above, the nozzle segment assembly 102 may also
include an inner ring support segment 114 and an outer ring support
segment 116. In several embodiments, the strut vane 106 and the
endwalls 108, 110 of the nozzle fairing 104, along with the
secondary vanes 112, may be configured to be sandwiched between the
inner and outer ring support segments 114, 116. For example, as
shown in the illustrated embodiment, the inner ring support segment
114 may generally correspond to the radially innermost portion of
the nozzle segment assembly 102 and may be configured to be
positioned directly adjacent to the inner endwall 108 when the
various components are assembled together. Similarly, the outer
ring support segment 116 may generally correspond to the radially
outermost portion of the nozzle segment assembly 102 and may be
configured to be positioned directly adjacent to the outer endwall
110 when the various components are assembled together.
[0034] In several embodiments, the inner ring support segment 114
may include an arcuate inner band segment 160 configured to be
positioned directly adjacent to the inner endwall 108 such that the
inner vane ends 130 of the secondary vanes 112 may be received
within and/or coupled to portions of the inner band segment 160.
For instance, as shown in FIG. 4, first and second vane recesses
162, 164 may be defined in an outer surface 166 of the inner band
segment 160 for receiving the inner vane ends 130 of the first and
second secondary vanes 112, respectively. Specifically, as
indicated above, the secondary vanes 112 may be configured to
extend through the inner endwall 108 such that the inner vane end
130 of each vane 112 projects radially inwardly from the inner
endwall 108. In such an embodiment, the inner vane ends 130 of the
secondary vanes 112 may be received within the corresponding vane
recesses 162, 164 defined by the inner band segment 160 for
supporting the secondary vanes 112 within the assembly 102.
[0035] It should be appreciated that, in several embodiments, the
inner vane ends 130 of the secondary vanes 112 may be configured to
be axially and/or circumferentially constrained within the vane
recesses 162, 164 while still being allowed to freely move radially
relative to the inner band segment 102. As shown in FIG. 4, in one
embodiment, such axial and/or circumferential constraint of the
inner vane ends 130 of the secondary vanes 112 may be accomplished
via a tab 198 extending radially outwardly from the inner vane end
130 of each secondary vane 112.
[0036] Additionally, in several embodiments, the inner ring support
segment may include or be coupled to an inter-stage sealing
mechanism 168 (shown in dashed lines) extending outwardly from the
inner band segment 160. In such embodiments, the inter-stage
sealing mechanism 168 may be configured to have any suitable
configuration known in the art for allowing the mechanism 168 to
provide sealing between adjacent rotor disks and/or between a rotor
disk and a separate structure of the gas turbine engine 10. For
example, the inter-stage sealing mechanism 168 may include one or
more honeycomb elements for sealing against corresponding rotor
teeth or other projections extending outwardly from the adjacent
rotor disk(s) of the gas turbine engine 10.
[0037] As shown in the illustrated embodiment, the outer ring
support segment 116 of the nozzle segment assembly 102 may
generally include an arcuate outer band segment 170 configured to
be positioned directly adjacent to the outer endwall 110 such that
the outer vane ends 132 of the secondary vanes 112 may be received
within and/or coupled to portions of the outer band segment 170.
For instance, in several embodiments, first and second vane
recesses 172, 174 may be defined in an inner surface (not shown) of
the outer band segment 170 for receiving the outer vane ends 132 of
the first and second secondary vanes 112, respectively.
Specifically, as indicated above, the secondary vanes 112 may be
configured to extend through the outer endwall 110 such that the
outer vane end 132 of each secondary vane 112 projects radially
outwardly from the outer endwall 110. In such an embodiment, the
outer vane ends 132 of the first and second secondary vanes 112 may
be received within the corresponding vane recesses 172, 174 defined
by the outer band segment 170 for supporting the secondary vanes
112 within the assembly 102.
[0038] It should be appreciated that, in several embodiments, the
outer vane ends 132 of the secondary vanes 112 may be configured to
be secured within the vane recesses 172, 174 using any suitable
attachment means and/or method. For example, as shown in the
illustrated embodiment, each outer vane end 132 may include a
mounting tab 176 configured to project radially outwardly
therefrom. As particularly shown in FIG. 4, each mounting tab 176
may define a through-hole 178 configured to receive a mounting pin
180. In such an embodiment, the recesses 172, 174 defined in the
outer band segment 170 may be shaped or otherwise formed so as to
accommodate mounting the secondary vanes 112 to the outer ring
support 116 through a pinned connection via the mounting pins 180.
For example, as shown in FIGS. 3 and 4, each recess 172, 174 may
include a raised portion 182 configured to receive the mounting tab
176 of the corresponding vane 112. In addition, a through-hole 184
may be defined through the raised portion 182 of each recess 172,
174 that is configured to be aligned with the through-hole 178
defined in each mounting tab 176 when the outer vane ends 132 of
the secondary vanes 112 are received within the recesses 172, 174.
As such, a mounting pin 180 may be inserted through the aligned
through-holes 178, 184 to secure each secondary vane 112 to the
outer ring support segment 116. In alternative embodiments, it
should be appreciated that the outer vane ends 132 of the secondary
vanes 112 may be secured within the recesses 172, 174 using any
other suitable attachment means, such as by bolting the ends 132
within each recess 172, 175.
[0039] It should also be appreciated that the outer band segment
170 may also include suitable features for mounting and/or sealing
the outer ring support segment 116 to a static structure (not
shown) within the gas turbine engine 10. For example, as shown in
FIG. 4, both a forward end 186 and an aft end 188 of the outer band
segment 170 may include features, such as slots, grooves, tabs
and/or the like, for mounting/sealing the outer ring support
segment 116 within the gas turbine engine 10.
[0040] Additionally, as particularly shown in FIG. 4, the outer
ring support segment 116 may also include a cantilevered strut 190
extending radially inwardly from the outer band segment 170.
Specifically, in several embodiments, the strut 190 may be
configured to extend radially inwardly from the outer band segment
170 though the hollow interior of the strut vane 106 to the inner
ring support segment 114. As a result, the strut 190 may be
configured to serve as the primary load-bearing component for the
nozzle fairing 104. Specifically, mechanical loads applied through
the nozzle fairing 104 may be transferred to the strut 190 and
carried radially outwardly to the outer band segment 170, at which
point the loads may be transferred to the static structure to which
the outer ring support segment 116 is coupled.
[0041] As shown in FIG. 4, to allow the strut 190 to be received
within the strut vane 106, strut openings 191 may be defined
through the inner and outer endwalls 108, 110 (only the strut
opening 191 defined through the outer endwall 110 being shown) that
are aligned with the interior cavity of the strut vane 106. Thus,
when assembling the various components of the disclosed assembly
102, the strut 190 may be inserted into the interior cavity of the
strut vane 106 via the strut opening 191 defined in the outer
endwall 10. In addition, a strut recess 192 may be defined in the
inner band segment 160 of the inner ring support segment 114 that
is configured to receive a free or tip end 193 of the strut 190.
Specifically, in several embodiments, when the strut 190 is
inserted fully within the nozzle fairing 104 (e.g., such that the
outer band segment 170 contacts or is otherwise positioned directly
adjacent to the outer endwall 110), the tip end 193 of the strut
190 may project radially inwardly from the inner endwall 108. In
such embodiments, the tip end 193 may be received within the strut
recess 192 defined by the inner band segment 160 to allow the strut
190 to be coupled between the outer and inner ring support segments
114, 116. For example, the tip end 193 of the strut 190 may be
secured within the strut recess 192 using a pinned connection, by
brazing or welding the tip end 193 within the recess 192 and/or by
using any other suitable attachment means and/or method.
[0042] It should be appreciated that, in several embodiments, the
strut 190 and the outer band segment 170 of the outer ring support
116 may be configured to be formed as a single, unitary component.
In such embodiments, the strut 190 and the outer band segment 170
may be formed integrally using any suitable manufacturing process
known in the art, such as casting, molding, machining, etc.
However, in an alternative embodiment, the strut 190 and the outer
band segment 170 may be manufactured as separate components. In
such an embodiment, the strut 190 may be configured to be coupled
to the outer band segment 170 using any suitable means and/or
method that allows the strut to be cantilevered from the outer band
segment 170.
[0043] It should also be appreciated that, in several embodiments,
both the strut 190 and the outer band segment 170 may be formed
from a metallic material. For example, suitable metallic materials
may include, but are not limited to, single-crystal and
non-single-crystal nickel alloys.
[0044] Additionally, in several embodiments, the strut 190 may
generally be configured to define an aerodynamic, airfoil-shaped
cross-section. Specifically, as shown in FIG. 5, the strut may
define a cross-sectional shape generally corresponding to the
aerodynamic cross-sectional shape of the strut vane 106. However,
in other embodiments, the strut 190 may define any other suitable
shape that allows it to be inserted within the strut vane 106.
[0045] Moreover, in several embodiments, one or more flow paths may
be defined around and/or in the strut 190 for receiving a cooling
medium within the strut vane 106 for cooling the vane 106 and/or
the strut 190. For example, as shown in FIG. 5, a gap 194 may be
defined between an inner surface 195 of the strut vane 106 and an
outer surface 196 of the strut 190 that defines a flow path for
receiving a cooling medium between the strut vane 106 and the strut
190. In addition, as shown in FIG. 5, the strut 190 may define one
or more internal cooling channels 197 for receiving a cooling
medium within the strut 197. In such an embodiment, the internal
cooling channels 197 may be fluidly isolated from the gap 194
defined between the strut vane 106 and the strut 190.
Alternatively, the internal cooling channels 197 may be in fluid
communication with the gap 194. For example, suitable cooling holes
(not shown) may be defined through the strut 190 between its outer
surface 196 and the cooling channel(s) 197 for directing a portion
of the cooling medium flowing through the cooling channel(s) 197
into the gap 194.
[0046] This written description uses examples to disclose the
invention, including the best mode, and also to enable any person
skilled in the art to practice the invention, including making and
using any devices or systems and performing any incorporated
methods. The patentable scope of the invention is defined by the
claims, and may include other examples that occur to those skilled
in the art. Such other examples are intended to be within the scope
of the claims if they include structural elements that do not
differ from the literal language of the claims, or if they include
equivalent structural elements with insubstantial differences from
the literal languages of the claims.
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